1. Field of the Invention
The present invention relates generally to rotary kinetic fluid motors or pumps, and more specifically to a turbine airfoil with end-wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is passed through a turbine to produce mechanical power. One method of increasing the efficiency of the gas turbine engine is to increase the temperature of the flow through the turbine. A typically turbine includes four stages of stationary vanes (also referred to as a nozzle) and rotor blades (also referred to as buckets) arranged in an alternating manner such that the vanes guide the flow into the blades. The first stage vane is exposed to the hottest temperature flow since the vane is located directly downstream from the combustor.
In order to allow for temperature higher than the melting temperature of the material used in the vane, designers have used complex cooling passages through the airfoils to provide cooling and therefore allow for higher flow temperatures to increase efficiency. Also, since the cooling air supplied to the airfoils for cooling must be under high pressures to prevent backflow from the hot gas into the airfoils, the cooling air is generally supplied from a middle stage of the compressor. Diverting compressed air from the compressor instead of using it with a fuel in the combustor also reduces the efficiency because the work used for compressing the cooling air is generally lost. Thus, providing an improvement in the cooling of the airfoil and reducing the amount of cooling air used for the same amount of cooling effectiveness would improve the efficiency of the engine. Higher efficiency means more power for the same amount of fuel.
U.S. Pat. No. 5,417,545 issued to Harrogate on May 23, 1995 entitled COOLED TURBINE NOZZLE ASSEMBLY AND METHOD OF CALCULATING THE DIAMETERS OF COOLING HOLES FOR USE IN SUCH AN ASSEMBLY discloses a turbine nozzle (vane) with an outer platform having an airfoil extending therefrom, and 2 rows of angled cooling holes located on the upstream end of the upper platform to supply cooling air to the platform and cooling the vane (see FIG. 1). The platform forms a smooth transition of the hot gas flow from the combustor to the guide vanes and is therefore exposed to the hot gas flow temperature. The cooling holes deliver necessary cooling to the transition platform.
As a result of the Harrogate cooling construction, stream-wise and circumferential cooling flow control due to airfoil external hot gas temperature and pressure variation is difficult to achieve. Film cooling air discharged from the double film rows has a tendency to migrate from the pressure side toward the vane suction surface which induce an uneven distribution of film cooling flow and end-wall metal temperature.
It is therefore an object of the present invention to provide for an improvement in the cooling of a leading edge end-wall (platform) of a turbine vane from that of the Harrogate patent.